Two spool gas turbine engine with interdigitated turbine section

ABSTRACT

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section that includes a first rotating component and a second rotating component. The first rotating component includes an inner shroud and an outer shroud outward of the inner shroud in the radial direction. The outer shroud defines a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction. The first rotating component further includes at least one connecting airfoil coupling the inner shroud and the outer shroud. The second rotating component is upstream of the one or more connecting airfoils of the first rotating component along the longitudinal direction. The second rotating component includes a plurality of second airfoils extended outward in the radial direction. The first rotating component defines at least one stage of the plurality of outer shroud airfoils upstream of the second rotating component.

FIELD

The present subject matter relates generally to gas turbine enginearchitecture. More particularly, the present subject matter relates to aturbine section for gas turbine engines.

BACKGROUND

Gas turbine engines generally include a turbine section downstream of acombustion section that is rotatable with a compressor section to rotateand operate the gas turbine engine to generate power, such as propulsivethrust. General gas turbine engine design criteria often includeconflicting criteria that must be balanced or compromised, includingincreasing fuel efficiency, operational efficiency, and/or power outputwhile maintaining or reducing weight, part count, and/or packaging (i.e.axial and/or radial dimensions of the engine).

Conventional gas turbine engines generally include turbine sectionsdefining a high pressure turbine in serial flow arrangement with anintermediate pressure turbine and/or low pressure turbine. The highpressure turbine includes an inlet or nozzle guide vane between thecombustion section and the high pressure turbine rotor. Conventionally,combustion gases exiting the combustion section define a relatively lowvelocity compared to a velocity (e.g., along a circumferential ortangential direction) of the first rotating stage of the turbine,generally defined as the high pressure turbine rotor. Thus,conventionally, the nozzle guide vane serves to accelerate a flow ofcombustion gases exiting the combustion section to more closely match orexceed the high pressure turbine rotor speed along a tangential orcircumferential direction. Such acceleration of flow using a nozzleguide vane to match or exceed high pressure turbine rotor speed is knownto improve general engine operability and performance.

Furthermore, conventional gas turbine engine turbine sections generallyinclude successive rows or stages of stationary and rotating airfoils,or vanes and blades, respectively. This conventional configurationgenerally conditions a flow of the combustion gases entering and exitingeach stage of vanes and blades. However, conventional turbine sections,and especially stationary airfoils (i.e. vanes and nozzle guide vanes)require considerable quantities and masses of cooling air to mitigatedamage due to hot combustion gases. For example, generally, nozzle guidevanes are designed to withstand a maximum combustion gas temperaturealong an annulus (i.e. hot spots), which may be significantly largerthan an average combustion gas temperature along the annulus. Thus,conventional engines are designed to use significant quantities ormasses of cooling air from a compressor section or unburned air from thecombustion section to mitigate structural damage, wear, deterioration,and ultimately, maintenance and repair, of the nozzle guide vanes.However, this cooling air adversely affects overall engine efficiency,performance, fuel consumption, and/or operability by removing energythat could otherwise be used in combustion to drive the turbines,compressors, and fan. Still further, the nozzle guide vane is often alimiting component when determining maintenance and repair intervals forgas turbine engines, thereby limiting overall engine performance andefficiency.

A known solution to improve turbine section efficiency is tointerdigitate the rotors of the turbine section (i.e. successive rows orstages of rotating airfoils or blades). For example, a known solution isto configure a turbine section, in serial flow arrangement from anupstream end to a downstream end along a longitudinal direction, with anozzle guide vane, a high pressure turbine rotor, another turbine vanestage (i.e. stationary airfoils), and an intermediate pressure turbineinterdigitated with a low pressure turbine. Another known solution is toconfigure a turbine section, in serial flow arrangement, with a nozzleguide vane, a high pressure turbine rotor, and various levels ofinterdigitated rotors thereafter, including low, intermediate, or highpressure turbine rotors.

However, despite various known solutions, there exists a need for anengine including a turbine section that may enable additional stages ofinterdigitation. Still further, despite various known solutions, thereexists a need for a turbine section that may further reduce cooling airconsumption, increase engine efficiency, performance, and/oroperability, and/or reduce part quantities, weight, and/or packaging(i.e. axial and/or radial dimensions).

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

The present disclosure is directed to a gas turbine engine defining alongitudinal direction, a radial direction, and a circumferentialdirection, and wherein the gas turbine engine defines an upstream endand a downstream end along the longitudinal direction. The gas turbineengine includes a turbine section that includes a first rotatingcomponent and a second rotating component. The first rotating componentincludes an inner shroud and an outer shroud outward of the inner shroudin the radial direction. The outer shroud defines a plurality of outershroud airfoils extended inward of the outer shroud along the radialdirection. The first rotating component further includes at least oneconnecting airfoil coupling the inner shroud and the outer shroud. Thesecond rotating component is upstream of the one or more connectingairfoils of the first rotating component along the longitudinaldirection. The second rotating component includes a plurality of secondairfoils extended outward in the radial direction. The first rotatingcomponent defines at least one stage of the plurality of outer shroudairfoils upstream of the second rotating component.

In one embodiment, the first rotating component defines a lean anglerelative to the axial centerline in which the plurality of outer shroudairfoils and/or the one or more connecting airfoils each define anobtuse lean angle approximately perpendicular to the axial centerline orextending toward the upstream end along the longitudinal direction frominward to outward along the radial direction.

In another embodiment, the first rotating component defines a lean anglerelative to the axial centerline, and wherein the plurality of outershroud airfoils and/or the one or more connecting airfoils each definean acute lean angle extending toward the downstream end along thelongitudinal direction from inward to outward along the radialdirection.

In yet another embodiment, the first rotating component and the secondrotating component are in interdigitation along the longitudinaldirection.

In various embodiments, the second rotating component defines a highspeed turbine and the first rotating component defines a low speedturbine.

In one embodiment, the gas turbine engine defines, in serial flowarrangement along the longitudinal direction from the upstream end tothe downstream end, the plurality of outer shroud airfoils of the firstrotating component, the plurality of second airfoils of the secondrotating component, and the one or more connecting airfoils of the firstrotating component.

In still various embodiments, the inner shroud of the first rotatingcomponent defines a plurality of inner shroud airfoils extended outwardalong the radial direction. In one embodiment, the inner shroud extendsfrom the connecting airfoil toward the downstream end.

In various embodiments, the gas turbine engine further includes acombustion section arranged in serial flow arrangement with the turbinesection. The combustion section, the first stage of the first rotatingcomponent, and the second rotating component are in serial flowarrangement along the longitudinal direction from the upstream end tothe downstream end. In one embodiment, the gas turbine engine defines,in serial flow arrangement along the longitudinal direction from theupstream end to the downstream end, the combustion section, the firststage of the first rotating component, the second rotating component,and then the first rotating component. In yet another embodiment, thegas turbine engine further includes a compressor section comprising ahigh pressure compressor and a fan assembly defining one or more stagesof a plurality of blades. The fan assembly, the compressor section, thecombustion section, and the turbine section are in serial flowarrangement along the longitudinal direction from the upstream end tothe downstream end. The first rotating component is connected androtatable with the fan assembly by a first shaft and the second rotatingcomponent is connected and rotatable with the high pressure compressorby a second shaft.

In still various embodiments, the plurality of outer shroud airfoils atthe first stage are coupled to an axially extended hub disposed inwardalong the radial direction of the plurality of outer shroud airfoils. Inone embodiment, the plurality of outer shroud airfoils at the firststage is further coupled to an arm extended generally inward along theradial direction, and wherein the arm is coupled to the axially extendedhub, and wherein the axially extended hub extends generally in thelongitudinal direction toward the upstream end of the engine. In anotherembodiment, the gas turbine engine further includes a first turbinebearing in which the second rotating component is further coupled to asecond shaft extended toward the upstream end, and the first turbinebearing is disposed along the radial direction between the second shaftand the axially extended hub of the first rotating component. In yetvarious embodiments, the first turbine bearing defines an air bearing, afoil bearing, a roller bearing, or a ball bearing.

In still various embodiments, the inner shroud of the first rotatingcomponent defines an inner shroud diameter and the outer shroud of thefirst rotating component defines an outer shroud diameter, and whereinthe inner shroud diameter is approximately 115% or less of the outershroud diameter. In one embodiment, the inner shroud diameter isapproximately equal to the outer shroud diameter.

In various embodiments, the first rotating component defines betweenabout 3 and 10 stages inclusively. In one embodiment, the first rotatingcomponent defines at least two stages of the plurality of outer shroudairfoils upstream of the connecting airfoil.

In another embodiment, the first rotating component rotates in a firstdirection and the second rotating component rotates in a seconddirection opposite of the first direction.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross sectional view of an exemplary gas turbineengine incorporating an exemplary embodiment of a turbine sectionaccording to an aspect of the present disclosure;

FIG. 2 is the schematic cross sectional view of the exemplary gasturbine engine of FIG. 1, further including a reduction gearbox in thefan assembly;

FIG. 3 is a schematic cross sectional view of an embodiment of theturbine section shown in FIG. 1;

FIG. 4 is a schematic cross sectional view of another embodiment of theturbine section shown in FIG. 1; and

FIG. 5 is cross sectional view depicting exemplary blade pitch angles.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “low”, “intermediate”, “high”, or their respective comparativedegrees (e.g. -er, where applicable) each refer to relative speedswithin an engine unless otherwise specified. For example, a “lowturbine” or “low speed turbine” defines a rotational speed lower than a“high turbine” or “high speed turbine”. Alternatively, unless otherwisespecified, the aforementioned terms may be understood in theirsuperlative degree. For example, a “low turbine” may refer to the lowestrotational speed turbine within a turbine section, and a “high turbine”may refer to the highest rotational speed turbine within the turbinesection.

Embodiments of a gas turbine engine with an interdigitated turbinesection are generally provided. The interdigitated turbine sectionincludes a first rotating component extended in a longitudinaldirection, in which the first rotating component includes an innershroud, an outer shroud, and at least one connecting airfoil couplingthe inner shroud to the outer shroud. The outer shroud includes aplurality of airfoils extended inward along a radial direction. Theinterdigitated turbine section may include a second rotating component.The second rotating component may include a plurality of second airfoilsextended outward in the radial direction, in which the second rotatingcomponent is disposed between the plurality of airfoils of the firstrotating component, and in which at least one stage of the plurality ofouter shroud airfoils is forward or upstream of the second rotatingcomponent.

The embodiments of the gas turbine engine with interdigitated turbinesection shown and described herein may enable additional stages ofinterdigitation of the first rotating component and the second rotatingcomponent, up to and including one or more stages of the first rotatingcomponent forward or upstream of the second rotating component. Invarious embodiments, the gas turbine engine with interdigitated turbinesection may further reduce cooling air consumption, increase engineefficiency, performance, and/or operability, and/or reduce partquantities, weight, and/or packaging (i.e. axial and/or radialdimensions). Still further, the interdigitated turbine section mayreduce a product of axial flow area and the square of the rotationalspeed (the product referred to as “AN²”) relative to an engineincorporating a reduction gearbox, while additionally reducing anaverage work factor per stage of the turbine section.

Referring now to the drawings, FIG. 1 is a schematic cross sectionalview of an exemplary gas turbine engine 10 (herein referred to as“engine 10”), shown as a high bypass turbofan engine, incorporating anexemplary embodiment of a turbine section 90 according to an aspect ofthe present disclosure. Although further described below with referenceto a turbofan engine, the present disclosure is also applicable toturbomachinery in general, including propfan, turbojet, turboprop, andturboshaft gas turbine engines, including marine and industrial turbineengines and auxiliary power units. As shown in FIG. 1, the engine 10 hasa longitudinal or axial centerline axis 12 that extends there throughfor reference purposes. The engine 10 defines a longitudinal directionL, a radial direction R, and an upstream end 99 and a downstream end 98along the longitudinal direction L.

In general, the engine 10 may include a substantially tubular outercasing 18 that defines an annular inlet 20. The outer casing 18 encasesor at least partially flows, in serial flow arrangement, a compressorsection 21, a combustion section 26, and an interdigitated turbinesection 90 (herein referred to as “turbine section 90”). Generally, theengine 10 defines, in serial flow arrangement from the upstream end 99to the downstream end 98, the fan assembly 14, the compressor section21, the combustion section 26, and the turbine section 90. In theembodiment shown in FIG. 1, the compressor section 21 defines a highpressure compressor (HPC) 24. In other embodiments, the fan assembly 14may further include or define one or more stages of a plurality of fanblades 42 that are coupled to and extend outwardly in the radialdirection R from a fan rotor 15 and/or a first shaft 36. In variousembodiments, multiple stages of the plurality of fan blades 42 coupledto the first shaft 36 may be referred to as a low pressure compressor(LPC) 22.

An annular fan casing or nacelle 44 circumferentially surrounds at leasta portion of the fan assembly 14 and/or at least a portion of the outercasing 18. In one embodiment, the nacelle 44 may be supported relativeto the outer casing 18 by a plurality of circumferentially-spaced outletguide vanes or struts 46. At least a portion of the nacelle 44 mayextend over an outer portion (in radial direction R) of the outer casing18 so as to define a bypass airflow passage 48 therebetween.

Referring now to FIG. 2, a schematic cross sectional side view of theengine 10 shown in FIG. 1 further including a reduction gearbox 45 inthe fan assembly 14 is generally provided. The reduction gearbox 45 mayinclude an epicyclical gear train including a star gear and a pluralityof planet gears. The plurality of planet gears may each be fixed suchthat each planet gear rotates on a fixed axis relative to the star gear.An annular gear surrounds the plurality of planet gears and rotates andtransfers power and torque from the star gear through the plurality ofplanet gears. In various embodiments, the gearbox may further includeadditional planet gears disposed radially between the plurality ofplanet gears and the star gear, or between the plurality of planet gearsand the annular gear.

Referring now to FIGS. 3-4, an exemplary embodiment of the turbinesection 90 of the engine 10 is generally provided. The turbine section90 includes a first rotating component 110 extended along thelongitudinal direction L. The first rotating component 110 includes aninner shroud 112, an outer shroud 114, and at least one connectingairfoil 116 coupling the inner shroud 112 to the outer shroud 114. Theouter shroud 114 includes a plurality of outer shroud airfoils 118extended inward along the radial direction R. In various embodiments,the inner shroud 112 may include a plurality of inner shroud airfoils119 extended outward along the radial direction R.

The inner shroud 112 and the outer shroud 114 each extend generallyalong the longitudinal direction L. The inner shroud 112 and/or theouter shroud 114 may each extend at least partially in the radialdirection R. In various embodiments, the inner shroud 112 extends fromthe connecting airfoil 116. In one embodiment, the inner shroud 112further extends toward the downstream end 98 along the longitudinaldirection L. In still various embodiments, the outer shroud 114 extendsfrom the connecting airfoil 116 toward the upstream end 99 along thelongitudinal direction L toward the combustion section 26.

Referring still to FIGS. 3-4, the turbine section 90 may further includea second rotating component 120 disposed forward or upstream 99 of theone or more connecting airfoils 116 of the first rotating component 110.The second rotating component 120 includes a plurality of secondairfoils 122 extended outward along the radial direction R.

In various embodiments, the first rotating component 110 defines aplurality of stages of rotating airfoils, such as the plurality of outershroud airfoils 118 disposed along the longitudinal direction L, or theone or more connecting airfoils 116, or the plurality of inner shroudairfoils 119 disposed along the longitudinal direction L. In oneembodiment, the first rotating component 110 defines at least one stageforward or upstream 99 of the second rotating component 120. In anotherembodiment, the turbine section 90 defines a first stage of airfoils inwhich the first stage includes the plurality of outer shroud airfoils118 of the first rotating component 110 forward or upstream 99 of eachstage of the second rotating component 120.

In various embodiments, such as shown in FIGS. 3-4, the engine 10defines, in serial flow arrangement along the longitudinal direction Lfrom the upstream end 99 to the downstream end 98, the plurality ofouter shroud airfoils 118 of the first rotating component 110, theplurality of second airfoils 122 of the second rotating component 120,and the one or more connecting airfoils 116 of the first rotatingcomponent 110. In still various embodiments, additional iterations ofinterdigitation between the first rotating component 110 and the secondrotating component 120 may be defined forward or upstream 99 of theconnecting airfoils 116.

Referring still to FIGS. 3-4, the engine 10 defines, in serial flowarrangement along the longitudinal direction L from the upstream end 99to the downstream end 98, the combustion section 26 and the turbinesection 90. More specifically, the engine 10 may define the serial flowarrangement of the combustion section 26, the first stage 101 of thefirst rotating component 110, and the second rotating component 120. Invarious embodiments, additional iterations of interdigitation betweenthe first rotating component 110 and the second rotating component 120may be defined aft or downstream 98 of the aforementioned arrangement.

For example, as shown in FIGS. 3-4, the engine 10 may further define theserial flow arrangement of the first rotating component 110, the secondrotating component 120, and the first rotating component 110. As anothernon-limiting example, the engine 10 may further define the serial flowarrangement of the plurality of outer shroud airfoils 118, the pluralityof second airfoils 122, the plurality of outer shroud airfoils 118, theplurality of second airfoils 122, and the connecting airfoils 116. Itshould be appreciated that although FIGS. 3-4 shows the second rotatingcomponent 120 as defining two stages, the second rotating component 120may define generally one or more stages between the first stage 101 ofthe first rotating component 110 and the connecting airfoils 116 of thefirst rotating component 110, and interdigitated therebetween along thelongitudinal direction L.

Referring now to FIGS. 1-4, in various embodiments, the first rotatingcomponent 110 defines a low speed turbine 30 drivingly connected androtatable with a first shaft 36 defining low speed shaft. In oneembodiment, as shown in FIG. 1, the first shaft 36 is connected to thefan assembly 14, of which is driven in rotation by the first rotatingcomponent 110 of the turbine section 90. The first shaft 36 is connectedto the fan rotor 15 of the fan assembly 14. In another embodiment, asshown in FIG. 2, the first shaft 36 is connected to the reductiongearbox 45 from the downstream end 98 and the fan rotor 15 is connectedto the reduction gearbox 45 from the upstream end 99. In variousembodiments, as shown in FIGS. 1 and 2, the fan assembly 14 defines aplurality of stages of the plurality of fan blades 42, of which furtherdefine the LPC 22. The LPC 22 is connected to and rotatable with thefirst shaft 36 and the plurality of fan blades 42 is connected androtatable with the fan rotor 15. In the embodiment shown in FIG. 2, thereduction gearbox 45 is disposed between the LPC 22 and the plurality offan blades 42 such that LPC 22, rotatable with the first shaft 36, isrotating at a first speed and the fan rotor 15, to which the pluralityof fan blades 42 is connected, is rotating at a proportionally reducedsecond speed relative to the first speed.

Referring still to FIGS. 1-4, the second rotating component 120 of theturbine section 90 defines a high speed turbine 28 drivingly connectedand rotatable with a second shaft 34 defining a high speed shaft. Thesecond shaft 34 is connected to the HPC 24, of which is driven inrotation by the second rotating component 120 of the turbine section 90.In various embodiments, the second rotating component 120 defining thehigh speed turbine 28 rotates generally at a higher rotational speedthan the first rotating component 110 defining the low speed turbine 30.

During operation of the engine 10, as shown in FIGS. 1-5 collectively, avolume of air as indicated schematically by arrows 74 enters the engine10 through an associated inlet 76 of the nacelle and/or fan assembly 14.As the air 74 passes across the fan blades 42, a portion of the air asindicated schematically by arrows 78 is directed or routed into thebypass airflow passage 48 while another portion of the air as indicatedschematically by arrows 80 is directed or through the fan assembly 14.Air 80 is progressively compressed as it flows through the compressorsection 21 toward the combustion section 26.

The now compressed air, as indicated schematically by arrows 82, flowsinto the combustion section 26 where a fuel 91 is introduced, mixed withat least a portion of the compressed air 82, and ignited to formcombustion gases 86. The combustion gases 86 flow into the turbinesection 90, causing rotary members of the turbine section 90 to rotateand support operation of respectively coupled rotary members in thecompressor section 21 and/or fan assembly 14.

In various embodiments, the first rotating component 110, and the firstshaft 36 to which it is attached, rotates in a first direction 161(shown in FIG. 5) along the circumferential direction C. The secondrotating component 120, and the second shaft 34 to which it is attached,rotates in a second direction 162 (shown in FIG. 5) opposite of thefirst direction 161 along the circumferential direction C. Althoughfurther described herein as a counter-rotating turbine engine, in whichthe first rotating component 110 rotates in a direction opposite of thesecond rotating component 120, it should be understood that thestructures provided herein enable the engine 10 to be configured as aco-rotating engine, in which the first rotating component 110 and thesecond rotating component 120 each rotate in the first direction 161.

It should further be understood that the first direction 161 and thesecond direction 162 as used and described herein are intended to denotedirections relative to one another. Therefore, the first direction 161may refer to a clockwise rotation (viewed from downstream lookingupstream) and the second direction 162 may refer to a counter-clockwiserotation (viewed from downstream looking upstream). Alternatively, thefirst direction 161 may refer to a counter-clockwise rotation (viewedfrom downstream looking upstream) and the second direction 162 may referto a clockwise rotation (viewed from downstream looking upstream).

Still further during an operation of the engine 10, combustion gases 86exiting the combustion section 26 define a generally low speed towardthe downstream end 98 of the engine 10. A low speed rotation (e.g. alonga tangential or circumferential direction C, as shown in FIG. 5) of thefirst stage 101 of the first rotating component 110 accelerates a speedof the combustion gases 86, such as in the tangential or circumferentialdirection C (shown in FIG. 5), to approximately equal or greater than aspeed of the second rotating component 120.

By defining the first rotating component 110 as the first stage 101 ofthe turbine section 90 aft or downstream of the combustion section 26,the engine 10 may obviate the need for a first turbine vane or nozzleguide vane to accelerate the combustion gases 86 forward or upstream ofthe second rotating component 120 defining a high speed turbine. Assuch, the engine 10 may reduce a quantity or mass of cooling air fromthe compressor section 21 and/or combustion section 26, therebyincreasing engine efficiency by enabling more energy (i.e. compressedair) to be used during combustion. Additionally, or alternatively, theturbine section 90 may reduce necessary cooling air and enable increasedperformance and/or operability of the compressor section 21, includingsurge margin and/or efficiency, or decrease a required amount of workfrom the compressor section 21, which may reduce axial dimensions orstages of the compressor section 21 and further reduce engine packaging,weight, and/or part count, and generally improve engine 10 performance.

Additionally, obviating the need for the first turbine vane or nozzleguide vane may enable the turbine section 90, or more specifically, thefirst stage 101, as a rotating stage, to be designed to an averagecombustion gas 86 temperature rather than designed to accommodate peaktemperatures (i.e. high spots) along an annulus of the core flowpath 70within the combustion section 26. Therefore, as all of the plurality ofouter shroud airfoils 118 of the first stage 101 are rotating, all ofthe plurality of outer shroud airfoils 118 may only transiently endureadverse effects of combustion hot spots rather than substantiallysustained or constant exposure to a higher temperature from thecombustion gases in contrast to other locations about the annulus of thecore flowpath 70. Still further, the turbine section 90 described hereinmay enable alternative design methods for the combustion section 26 dueto a decreased adverse effect of combustion hot spots on the turbinesection 90. Therefore, the turbine section 90 may enable design of thecombustion section 26 to further improve combustion stability, decreaseemissions, increase operability across all or part of a flight envelope,increase altitude re-light performance, and/or decrease lean blowout(LBO).

Referring now to FIG. 5, exemplary embodiments of orientations ofairfoils 170 of the first and second rotating components 110, 120 aregenerally provided. FIG. 5 generally depicts angular orientations andprofiles of various embodiments of the airfoils 170, in which theairfoils 170 may be representative of the plurality of outer shroudairfoils 118, the plurality of second airfoils 122, the one or moreconnecting airfoil(s) 116, or the plurality of inner shroud airfoils119. The airfoils 170 depicted in FIG. 5 generally describe, at least inpart, aerodynamic structures inducing the first direction 161 ofrotation along the circumferential direction C or the second rotation162 opposite of the first direction 161 for the first and/or secondrotating components 110, 120.

The airfoils 170 may be arranged along the circumferential direction Cinto a plurality of stages 171, 172 separated along the longitudinaldirection L. The first direction stage 171 shown in FIG. 5 maycorrespond to the first stage 101 of the first rotating component 110shown in FIGS. 3-4. Still further, orientations and/or profiles of theairfoils 170 of the first direction stage 171 may correspond generallyto the airfoils of the first rotating component 110, such as theplurality of outer shroud airfoils 118, the connecting airfoils 116, andthe plurality of inner shroud airfoils 119. The airfoils 170 shown inthe second direction stage 172 shown in FIG. 5 may correspond to theplurality of second airfoils 122 of the second rotating component 120shown in FIGS. 3-4.

In various embodiments, the airfoil 170 may define a first exit angle178 defined by an angular relationship of the axial centerline 12 to anexit direction 177 of the combustion gases 86 passing the airfoil 170along the longitudinal direction L from the upstream end 99 toward thedownstream end 98. The resulting first exit angle 178 may define theairfoil 170 such that the flow of combustion gases 86 across eachairfoil 170 from the upstream end 99 toward the downstream end 98induces the first direction 161 of rotation in the circumferentialdirection C.

In other embodiments, the airfoil 170 may define a second exit angle 179defined by an angular relationship to the axial centerline 12 to theexit direction 177 of the combustion gases 86, in which the exitdirection 177 extends generally opposite for the second exit angle 179relative to the first exit angle 178. The resulting second exit angle179 may define the airfoil 170 such that the flow of combustion gases 86across each airfoil 170 induces the second direction 162 of rotation inthe circumferential direction C.

It should be appreciated that the first exit angle 178 and the secondexit angle 179 each define general angular relationships relative theaxial centerline 12, such as a positive or negative acute angle.Therefore, each airfoil 170 defining the first exit angle 178 (or,alternatively, the second exit angle 179) may define differentmagnitudes of angles at each stage of airfoils, in which each angledefines a generally positive acute angle relative to the axialcenterline 12 (or, alternatively, a generally negative acute angle forthe second exit angle 179).

Referring still to FIG. 5, the airfoil 170 may define a suction side 173and a pressure side 174. The first direction stage 171 may define thesuction side 173 as convex toward the first direction 161 and thepressure side 174 as concave toward the first direction 161 such thatthe airfoils 170 rotate in the first direction 161. The second directionstage 172 may define the suction side 173 as convex toward the seconddirection 162 opposite of the first direction 161 and the pressure side174 as concave toward the second direction 162 such that the airfoils170 rotate in the second direction 162.

Referring to FIGS. 1-5, in one embodiment, the airfoils 170 of the firstrotating component 110 are generally configured as defined in the firstdirection stage 171 to rotate in the first direction 161. The airfoils170 of the second rotating component 120 are generally configured as thesecond direction stage 172 to rotate in the second direction 162opposite of the first direction 161. For example, the first rotatingcomponent 110, as the low speed turbine 30 coupled to the first shaft36, may rotate clockwise viewed from the downstream end 98 lookingtoward the upstream end 99. The second rotating component 120, as thehigh speed turbine 28 coupled to the second shaft 34, may rotatecounter-clockwise when viewed from the downstream end 98 toward theupstream end 99.

Each stage of the first rotating component 110 and the second rotatingcomponent 120 defines a plurality of rotating airfoils 170 (e.g. blades,shown in FIG. 5). Although FIGS. 3-4 depicts the turbine section 90 asdefining two stages forward of the connecting airfoils 116 of the firstrotating component 110 interdigitated with the second rotating component120 defining two stages, it should be understood that the turbinesection 90 may include other quantities or combinations ofinterdigitation. For example, the first and second rotating components110, 120 may together define additional iterations of stages disposedupstream 99 or forward of the connecting airfoil(s) 116. As anothernon-limiting example, the second rotating component 120 may define asingle stage interdigitated among stages of the first rotating component110.

In various embodiments of the engine 10 including the turbine section 90shown in FIGS. 1-5, the first rotating component 110 defines betweenabout 3 and 10 stages (inclusively) i.e. between about 3 and 10 rows ofpluralities of rotating airfoils separated along the longitudinaldirection L. In one embodiment, the first rotating component 110 definesat least one stage upstream or forward of the second rotating component120 (e.g. the first stage 101). In another embodiment, the firstrotating component 110 defines at least two stages of the plurality ofouter shroud airfoils 118 upstream of the connecting airfoil 116. In yetanother embodiment, the first rotating component 110 defines at leasttwo stages downstream or aft of the second rotating component 120 (e.g.the connecting airfoil 116, the plurality of outer shroud airfoils 118,and/or the plurality of inner shroud airfoils 119, or combinationsthereof).

Referring back to FIGS. 3-4, exemplary embodiments of the engine 10 aregenerally provided, in which one or more stages of the plurality ofouter shroud airfoils 118 and/or the connecting airfoils 116 define alean angle 109 relative to the axial centerline 12 and the radialdirection R. As discussed herein, the lean angle 109 is definedcounterclockwise from the axial centerline 12 from the downstream end 98toward the upstream end 99.

In the embodiment shown in FIG. 3, the connecting airfoils 116 and/orthe outer shroud airfoils 118 may each define an obtuse or forward leanangle 109 in which one or more of the connecting airfoils 116 and/or theouter shroud airfoils 118 extend toward the downstream end 98 (i.e.radially inward ends of the airfoils 116, 118 are further downstreamthan radially outward ends). The obtuse or forward lean angle 109 maycounteract or offset centrifugal loads on the plurality of outer shroudairfoils 118 and/or connecting airfoils 116 during rotation of theturbine section 90. The obtuse lean angle 109 may also, oralternatively, counteract or offset axial loads on each of the airfoils116, 118 during operation of the engine 10, such as due to rotationand/or the flow of combustion gases 86 through a core flowpath 70. Stillfurther, the obtuse or forward lean angle 109 may dispose the airfoils116, 118 generally perpendicular to the core flowpath 70 downstream ofeach airfoil 116, 118.

However, in the embodiment shown in FIG. 4, the connecting airfoils 116and/or the outer shroud airfoils 118 may each define a generallyperpendicular or acute lean angle 109 in which one or more of theconnecting airfoils 116 and/or the outer shroud airfoils 118 extendgenerally radially outward from the axial centerline 12 or toward theupstream end 99 (i.e. radially inward ends of the airfoils 116, 118 areapproximately equal to or further upstream than radially outward ends).

Referring still to FIGS. 3-4, the inner shroud 112 may define a maximuminner shroud diameter 107 and the outer shroud 114 may define an outershroud diameter 108. In one embodiment, the inner shroud diameter 107may be approximately equal to the outer shroud diameter 108. Forexample, the inner shroud diameter 107 defined at a last stage of thefirst rotating component 110 may be approximately equal to the outershroud diameter 108 at approximately the first stage 101. In otherembodiments, the inner shroud diameter 107 may be approximately 115% orless of the outer shroud diameter 108. In yet another embodiment, theinner shroud diameter 107 at the last stage of the first rotatingcomponent 110 may be approximately 115% or less of the outer shrouddiameter 108 at the first stage 101 of the first rotating component 110.In still other embodiments, the inner shroud diameter 107 may beapproximately 110% or less of the outer shroud diameter 108. In yetother embodiments, the inner shroud diameter 107 may be approximately105% or less of the outer shroud diameter 108.

The exemplary embodiment of the engine 10 shown in FIGS. 3-4 may furtherinclude a first turbine bearing 200 disposed radially within at leastthe combustion section 26 and/or the turbine section 90. In variousembodiments, the first turbine bearing 200 may define a generallynon-contacting air bearing or foil bearing. In various otherembodiments, the first turbine bearing 200 may define a generallycontacting bearing such as, but not limited to, a roller bearing or aball bearing. The first turbine bearing 200 may further enable theoverhung or cantilevered first rotating component 110 to extend forwardor upstream of the second rotating component 120.

In various embodiments, the second rotating component 120 is coupled tothe second shaft 34 extended toward the upstream end 99 of the engine10. The plurality of outer shroud airfoils 118 at the first stage 101may further be coupled to an axially extended hub 105 disposed inwardalong the radial direction R of the plurality of outer shroud airfoils118 at the first stage 101. In one embodiment, the plurality of outershroud airfoils 118 at the first stage 101 is further coupled to an arm106 extended generally inward along the radial direction R. The arm 106is coupled to the axially extended hub 105 in which the axially extendedhub 105 extends generally in the longitudinal direction L toward theupstream end 99. The first turbine bearing 200 is disposed between thesecond shaft 34 and the axially extended hub 105 of the first rotatingcomponent 110 along the radial direction R.

In one embodiment, the first turbine bearing 200 supports the firstrotating component 110 inward of the plurality of outer shroud airfoils118 at the first stage 101. For example, the first turbine bearing 200may support the overhung or cantilevered first rotating component 110generally forward or upstream 99 of the second rotating component 120.

In another embodiment, the first turbine bearing 200 supports the secondrotating component 120. In various embodiments, the first turbinebearing 200 supports the first rotating component 110 and the secondrotating component 120. For example, the first turbine bearing 200 maydefine a differential bearing disposed between the first rotatingcomponent 110 and the second rotating component 120 along the radialdirection R. In still various embodiments, the first turbine bearing 200may define an air bearing, a foil bearing, a roller bearing, or a ballbearing.

During operation of the engine 10, a flow of a lube, hydraulic, orpneumatic fluid (e.g. oil, air, etc.) may flow from the compressorsection 21 and/or through the combustion section 26 (e.g. along theradial direction R through one or more manifolds) to the first turbinebearing 200 to provide a protective film that may enable rotation andprotect the first rotating component 110, the second rotating component120, and the first turbine bearing 200 from damage due to friction,temperature, and other wear and degradation.

The arrangement of the first bearing 200 may provide support toward theupstream end 99 of the first rotating component 110 to be interdigitatedforward and/or among the second rotating component 120. Furthermore, thefirst bearing 200 provides support toward the upstream end 99 of thefirst rotating component 110 that limits an overhanging or cantileveredweight of the first rotating component 110 from the connecting airfoil116 upstream toward the combustion section 26. Still further, the firstbearing 200 provides support toward the upstream end 99 of the firstrotating component 110 that provides balance to the inner shroud 112 andthe plurality of inner shroud airfoils 119 extended therefrom toward thedownstream end 98 of the turbine section 90. In various embodiments, theaxially extended hub 105 of the first rotating component 110 may furtherdefine one or more balance planes. The balance plane may define featuresto which weight may be added to or removed from the first rotatingcomponent 110 to aid rotor balance and operation.

Referring still to FIGS. 2-3, the turbine section 90 further includesone or more turbine vanes 150. The turbine vane 150 may define aplurality of stationary airfoils (i.e. vanes) in circumferentialarrangement. In one embodiment, the turbine vane 150 is disposed betweenthe pluralities of inner shroud airfoils 119 along the longitudinaldirection L. In various embodiments, the turbine vane 150 is disposeddownstream 98 of the connecting airfoil 116 of the first rotatingcomponent 110. The turbine vane 105, or pluralities thereof,interdigitated among the pluralities of inner shroud airfoils 119 mayenable further conditioning of the combustion gases 86 and work orenergy extraction from the first rotating component 110 via theplurality of inner shroud airfoils 119.

The turbine section 90 shown and described herein may improve uponexisting turbine sections by providing improved fuel efficiency,operational efficiency, and/or power output while maintaining orreducing weight, part count, and/or packaging. The plurality of outershroud airfoils 118 of the first rotating component 110 interdigitatedamong the plurality of second airfoils 122 of the second rotatingcomponent 120 may reduce packaging and reduce part count by removingstages of stationary airfoils between each rotating component.Additionally, the turbine section 90 may provide efficiency benefitscomparable to a reduction gearbox without adding weight or size (e.g.axial length) to the engine 10. The first rotating component 110, as thefirst stage 101 downstream of the combustion section 26, may furtherimprove engine efficiency by reducing cooling air appropriated away fromproducing combustion gases 86, thereby allowing more energy from thecompressor section 21 to be used in combustion and operation of theengine 10. Furthermore, removing the nozzle guide vane between thecombustion section 26 and the first rotating component 110 of theturbine section 90 may reduce or eliminate design constraints related tohot spots in the combustion gases along the annulus of the core flowpath70.

The various embodiments of the turbine section 90 generally shown anddescribed herein may be constructed as individual blades installed intodrums or hubs, or integrally bladed rotors (IBRs) or bladed disks, orcombinations thereof. The blades, hubs, or bladed disks may be formed ofceramic matrix composite (CMC) materials and/or metals appropriate forgas turbine engine hot sections, such as, but not limited to,nickel-based alloys, cobalt-based alloys, iron-based alloys, ortitanium-based alloys, each of which may include, but are not limitedto, chromium, cobalt, tungsten, tantalum, molybdenum, and/or rhenium.The turbine section 90, or portions or combinations of portions thereof,including the inner shroud 112, the outer shroud 114, the connectingairfoil(s) 116, the plurality of outer shroud airfoils 118, and/or theplurality of inner shroud airfoils 119, may be formed using additivemanufacturing or 3D printing, or casting, forging, machining, orcastings formed of 3D printed molds, or combinations thereof. Theturbine section 90, or portions thereof, such as stages of the rotatingcomponents 110, 120, the outer shroud 114, the inner shroud 112, and/orvarious shrouds, seals, and other details may be mechanically joinedusing fasteners, such as nuts, bolts, screws, pins, or rivets, or usingjoining methods, such as welding, brazing, bonding, friction ordiffusion bonding, etc., or combinations of fasteners and/or joiningmethods. Still further, it should be understood that the first rotatingcomponent 110, including the inner and/or outer shroud 112, 114, mayincorporate features that allow for differential expansion. Suchfeatures include, but are not limited to, aforementioned methods ofmanufacture, various shrouds, seals, materials, and/or combinationsthereof

The systems shown in FIGS. 1-5 and described herein may decrease fuelconsumption, increase operability, increase engine performance and/orpower output while maintaining or reducing weight, part count, and/orpackaging (e.g. radial and/or axial dimensions). The systems providedherein may allow for increased bypass ratios and/or overall pressureratios over existing gas turbine engine configurations, such asturbofans, while maintaining or reducing packaging relative to other gasturbine engines of similar power output. The systems described hereinmay contribute to improved bypass ratio and/or overall pressure ratioand thereby increase overall gas turbine engine efficiency. The systemsprovided herein may increase overall gas turbine engine efficiency byreducing or eliminating stationary airfoils that require cooling air(e.g. nozzle guide vane). Additionally, the systems provided herein mayreduce gas turbine engine packaging and weight, thus increasingefficiency, by reducing rotating and/or stationary airfoil quantities(e.g. blades and/or vanes) by approximately 40% or more over gas turbineengines of similar size and/or power output.

Still further, the systems shown in FIGS. 1-5 and described herein mayreduce a product of a flow area and the square of the rotational speed(the product herein referred to as “AN²”) of the gas turbine engine. Forexample, engine 10 shown and described in regard to FIGS. 1-5 maygenerally reduce AN² relative to a conventional geared turbofanconfiguration. Generally, lowering the AN², such as by reducing therotational speed and/or the flow area, increases the required averagestage work factor (i.e. the average required loading on each stage ofrotating airfoils). However, the systems described herein may lower theAN² while also lowering the average stage work factor and maintainingaxial length of the turbine section 90 (compared to engines of similarthrust output and packaging) by interdigitating the first rotatingcomponent 110 among the one or more stages of the second rotatingcomponent 120 while also defining a non-digitated turbine structure(i.e. the inner shroud 112 and the plurality of inner shroud airfoils119) toward the downstream end 98 of the turbine section 90. Therefore,the first rotating component 110 may increase the quantity of rotatingstages of airfoils while reducing the average stage work factor, andtherefore the AN², while mitigating increases in axial length to producea similar AN² value. The first rotating component 110 may further reducethe AN² while additionally reducing the overall quantity of airfoils,rotating and stationary, in the turbine section 90 relative to turbinesections of gas turbine engines of similar power output and/orpackaging.

Furthermore, the systems shown in FIGS. 1-5 and described herein mayfurther improve engine efficiency, reduce airfoil quantity, reduceengine weight, and/or alleviate combustion section design constraints byinterdigitating the first rotating component 110 forward or upstream 99of the second rotating component 120 defining the high speed turbine 28.For example, defining the first stage of the first rotating component110 as immediately downstream 98 of the combustion section 26, without afirst turbine vane or nozzle guide vane therebetween, as well asdefining the first rotating component 110 in counter-rotation with thesecond rotating component 120, may reduce effects of overall combustionhot spots on the first stage of the first rotating component 110 incontrast to a stationary, first turbine vane or nozzle guide vane. Assuch, the turbine section 90 and engine 10 described herein may removeconstraints to combustion section 26 design by de-emphasizing hot spots,or combustion pattern factor, in favor of other design criteria, such asdecreasing emissions, improving lean blow-out (LBO) and/or altitudere-light, improving overall operability across part or all of anoperating envelope, or increasing the operating envelope.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A gas turbine engine, wherein the gas turbineengine defines a longitudinal direction, a radial direction, and acircumferential direction, and wherein the gas turbine engine defines anupstream end and a downstream end along the longitudinal direction, thegas turbine engine comprising: a turbine section comprising a firstrotating component and a second rotating component, wherein the firstrotating component includes an inner shroud and an outer shroud outwardof the inner shroud in the radial direction, wherein the outer shrouddefines a plurality of outer shroud airfoils extended inward of theouter shroud along the radial direction, and wherein the first rotatingcomponent further includes at least one connecting airfoil coupling theinner shroud and the outer shroud, and wherein the second rotatingcomponent is upstream of the one or more connecting airfoils of thefirst rotating component along the longitudinal direction, the secondrotating component includes a plurality of second airfoils extendedoutward in the radial direction, and wherein the first rotatingcomponent defines at least one stage of the plurality of outer shroudairfoils upstream of the second rotating component.
 2. The gas turbineengine of claim 1, wherein the first rotating component defines a leanangle relative to the axial centerline, and wherein the plurality ofouter shroud airfoils and/or the one or more connecting airfoils eachdefine an obtuse lean angle approximately perpendicular to the axialcenterline or extending toward the upstream end along the longitudinaldirection from inward to outward along the radial direction.
 3. The gasturbine engine of claim 1, wherein the first rotating component definesa lean angle relative to the axial centerline, and wherein the pluralityof outer shroud airfoils and/or the one or more connecting airfoils eachdefine an acute lean angle extending toward the downstream end along thelongitudinal direction from inward to outward along the radialdirection.
 4. The gas turbine engine of claim 1, wherein the firstrotating component and the second rotating component are ininterdigitation along the longitudinal direction.
 5. The gas turbineengine of claim 1, wherein the second rotating component defines a highspeed turbine and the first rotating component defines a low speedturbine.
 6. The gas turbine engine of claim 1, wherein the gas turbineengine defines, in serial flow arrangement along the longitudinaldirection from the upstream end to the downstream end, the plurality ofouter shroud airfoils of the first rotating component, the plurality ofsecond airfoils of the second rotating component, and the one or moreconnecting airfoils of the first rotating component.
 7. The gas turbineengine of claim 1, wherein the inner shroud of the first rotatingcomponent defines a plurality of inner shroud airfoils extended outwardalong the radial direction.
 8. The gas turbine engine of claim 7,wherein the inner shroud extends from the connecting airfoil toward thedownstream end.
 9. The gas turbine engine of claim 1, further comprisinga combustion section arranged in serial flow arrangement with theturbine section, and wherein the combustion section, the first stage ofthe first rotating component, and the second rotating component are inserial flow arrangement along the longitudinal direction from theupstream end to the downstream end.
 10. The gas turbine engine of claim9, wherein the gas turbine engine defines, in serial flow arrangementalong the longitudinal direction from the upstream end to the downstreamend, the combustion section, the first stage of the first rotatingcomponent, the second rotating component, and then the first rotatingcomponent.
 11. The gas turbine engine of claim 1, wherein the pluralityof outer shroud airfoils at the first stage are coupled to an axiallyextended hub disposed inward along the radial direction of the pluralityof outer shroud airfoils.
 12. The gas turbine engine of claim 11,wherein the plurality of outer shroud airfoils at the first stage isfurther coupled to an arm extended generally inward along the radialdirection, and wherein the arm is coupled to the axially extended hub,and wherein the axially extended hub extends generally in thelongitudinal direction toward the upstream end of the engine.
 13. Thegas turbine engine of claim 11, further comprising a first turbinebearing, and wherein the second rotating component is further coupled toa second shaft extended toward the upstream end, and wherein the firstturbine bearing is disposed along the radial direction between thesecond shaft and the axially extended hub of the first rotatingcomponent.
 14. The gas turbine engine of claim 13, wherein the firstturbine bearing defines an air bearing, a foil bearing, a rollerbearing, or a ball bearing.
 15. The gas turbine engine of claim 1,wherein the inner shroud of the first rotating component defines aninner shroud diameter and the outer shroud of the first rotatingcomponent defines an outer shroud diameter, and wherein the inner shrouddiameter is approximately 115% or less of the outer shroud diameter. 16.The gas turbine engine of claim 15, wherein the inner shroud diameter isapproximately equal to the outer shroud diameter.
 17. The gas turbineengine of claim 1, wherein the first rotating component defines betweenabout 3 and 10 stages inclusively.
 18. The gas turbine engine of claim17, wherein the first rotating component defines at least two stages ofthe plurality of outer shroud airfoils upstream of the connectingairfoil.
 19. The gas turbine engine of claim 1, wherein the firstrotating component rotates in a first direction and the second rotatingcomponent rotates in a second direction opposite of the first direction.20. The gas turbine engine of claim 9, further comprising: a compressorsection comprising a high pressure compressor; and a fan assemblydefining one or more stages of a plurality of blades, wherein the fanassembly, the compressor section, the combustion section, and theturbine section are in serial flow arrangement along the longitudinaldirection from the upstream end to the downstream end, and wherein thefirst rotating component is connected and rotatable with the fanassembly by a first shaft, and wherein the second rotating component isconnected and rotatable with the high pressure compressor by a secondshaft.